Heat exchanger

ABSTRACT

A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes a plurality of fan blades defining a fan diameter. The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and the heat exchanger module including a plurality of heat transfer elements for transfer of heat from a first fluid contained within the heat transfer elements to an airflow passing over a surface of the heat transfer elements prior to entry of the airflow into an inlet to the fan assembly. Each heat transfer element may be individually and independently fluidly isolated from the remaining heat transfer elements.

This disclosure claims the benefit of UK Patent Application No. GB2016007.3, filed on 9 Oct. 2020, which is hereby incorporated herein inits entirety.

FIELD OF THE DISCLOSURE

The present disclosure relates to a circumferential vane arrayaccommodating a heat exchanger and particularly to a circumferentialvane array accommodating a heat exchanger, for use with a turbofan gasturbine engine.

BACKGROUND TO THE DISCLOSURE

A conventional turbofan gas turbine engine uses heat exchangers to coola variety of fluids including inter alia air, fuel and oil. Typically,such heat exchangers use bypass air or an air offtake from thecompressor as the cooling medium. The heat exchanger itself may bepositioned in the bypass duct or externally to the engine with thecorresponding ducting.

The use of bypass air or a compressor offtake stream as the coolingmedium in a heat exchanger will adversely affect the performance of theengine, for example by reducing specific thrust or increasing specificfuel consumption. Alternatively, or additionally, such offtakes canadversely affect engine performance, for example by reducing surgemargin.

In a further alternative conventional arrangement, an air flow toprovide the cooling medium in a heat exchanger may be drawn separatelyfrom the air flow through the gas turbine engine. For example, in anairframe application the air flow providing the cooling medium may bedrawn from an air intake or duct separate from the engine.

As used herein, a range “from value X to value Y” or “between value Xand value Y”, or the likes, denotes an inclusive range; including thebounding values of X and Y. As used herein, the term “axial plane”denotes a plane extending along the length of an engine, parallel to andcontaining an axial centreline of the engine, and the term “radialplane” denotes a plane extending perpendicular to the axial centrelineof the engine, so including all radial lines at the axial position ofthe radial plane. Axial planes may also be referred to as longitudinalplanes, as they extend along the length of the engine. A radial distanceor an axial distance is therefore a distance in a radial or axial plane,respectively.

STATEMENTS OF DISCLOSURE

According to a first aspect of the present disclosure there is provideda turbofan gas turbine engine comprising, in axial flow sequence, a heatexchanger module, a fan assembly, a compressor module, a turbine module,and an exhaust module, the fan assembly comprising a plurality of fanblades defining a fan diameter (D), the heat exchanger module being influid communication with the fan assembly by an inlet duct, the heatexchanger module comprising a plurality of heat transfer elements fortransfer of heat from a first fluid contained within the heat transferelements to an airflow passing over a surface of the heat transferelements prior to entry of the airflow into an inlet to the fanassembly, wherein each heat transfer element may be individually andindependently fluidly isolated from the remaining heat transferelements.

In one arrangement of the present disclosure, the first fluid is an oilthat is used to provide cooling to other parts of the turbofan engine.The oil is circulated through the heat transfer elements while the airflow passing over the heat transfer elements enables the rejection ofheat energy from the oil to the air flow.

Any physical damage to the heat transfer element may result in leakageof the oil from the heat transfer element. Such physical damage might becaused to the heat transfer element by a foreign object carried throughon the vane air flow.

Leakage of oil resulting from damage is undesirable since the oil isrequired for the heat rejection process. If more than a pre-determinedquantity of oil is lost from the oil system, the turbofan engine maybecome inoperative.

An oil leakage arising from damage to a heat transfer element may resultin oil entering, for example, the secondary air system cavities. Thiscould pose a serious fire risk to the turbofan engine, which isundesirable. The secondary air system includes those air flows that arerequired for the safe operation of the turbofan engine such as interalia, the turbine cooling air, bearing and rim sealing air, bearingthrust control, active tip clearance control, and handling bleeds.

Additionally, the downstream air from the gas turbine engine, forexample a compressor bleed, may be used for ventilation purposes and thepresence of oil in such an air stream is undesirable.

The arrangement of the present disclosure enables a damaged heattransfer element to be isolated from the remaining heat transferelements. This will prevent further oil loss from the damaged elementand enable continued operation of the gas turbine engine albeit atperhaps slightly reduced heat energy rejection capability.

Optionally, each heat transfer element comprises an inlet and an outlet,and both the inlet and the outlet are selectably closable to fluidlyisolate the heat transfer element.

In one arrangement of the present disclosure, an actuatable valve ispositioned on each of the inlet to and outlet from each heat transferelement. In the event of an oil leak from the heat transfer element thevalves on each of the inlet to and outlet from the heat transfer elementcan be closed. This fluidly isolates the damaged heat transfer elementfrom the remainder of the heat transfer element.

In another arrangement, an actuatable valve may be provided on only oneof the inlet to and outlet from the heat transfer element. In thisarrangement, a one-way valve may be provided on the other of the inletto and outlet from the heat transfer element. In this way, if a leak ispresent in the heat transfer element then the actuatable valve may beclosed to fluidly isolate the heat transfer element and prevent furtherloss of oil from the system.

Optionally, the heat exchanger module further comprises a flow sensorconfigured to detect a flow rate of the first fluid through the heattransfer elements, and the heat exchanger module is configured toisolate a heat transfer element in the event that the detected flow rateis outside of a predetermined acceptable flow rate range.

The monitoring of the first fluid flow rate will enable a user todetermine whether there is any leakage of the first fluid. In responseto the leak determination, the respective heat transfer element can befluidly isolated from any other heat transfer elements.

In one arrangement, a flow sensor is provided on each heat transferelement. In an alternative arrangement, fewer flow sensors than heattransfer elements may be provided, for example at fluid junctionsbetween heat transfer elements.

Optionally, the heat exchanger module further comprises a pressuresensor configured to detect a pressure of the first fluid containedwithin the heat transfer elements, and the heat exchanger module isconfigured to isolate a heat transfer element in the event that thedetected pressure is outside of a predetermined acceptable pressurerange.

By monitoring the pressure in the first fluid as it is flowing itbecomes possible for a user to determine whether there is any leakage ofthe first fluid. The leaking heat transfer element can then be fluidlyisolated from any other heat transfer elements. In one example, a fluidpressure sensor is positioned on a fluid line to or from each heattransfer element.

Optionally, the heat exchanger module comprises a plurality ofradially-extending hollow vanes arranged in a circumferential array witha channel extending axially through the heat exchanger module betweeneach two circumferentially adjacent vanes, each of the hollow vanesaccommodating at least one heat transfer element.

In one arrangement of the present disclosure, each of the heat transferelements is enclosed within a hollow vane, with the hollow vanesarranged in a circumferential array. Positioning the heat transferelements inside a hollow vane provides a measure of physical protectionto the heat transfer elements. The hollow vane also provides for theattachment of ancillary components such as isolation valves and fluidsensors.

Optionally, the fan diameter D is within the range of 0.3 m to 2.0 m,preferably within the range 0.4 m to 1.5 m, and more preferably in therange of 0.7 m to 1.0 m.

In one embodiment of the disclosure, the fan diameter is 0.9 m.

Consequently, for the same heat energy loading rejected to the air flowthrough the heat exchanger, the loss in propulsive efficiency of theturbofan engine is proportionately smaller for a large diameter (forexample, approximately 1.5 to 2.0 m in diameter) turbofan engine thanfor a small diameter turbofan engine.

The fan tip radius, measured between a centreline of the engine and anoutermost tip of each fan blade at its leading edge, may be in the rangefrom 95 cm to 200 cm, for example in the range from 110 cm to 150 cm, oralternatively in the range from 155 cm to 200 cm. The fan tip radius maybe greater than any of: 110 cm, 115 cm, 120 cm, 125 cm, 130 cm, 135 cm,140 cm, 145 cm, 150 cm, 155 cm, 160 cm, 165 cm, 170 cm, 175 cm, 180 cm,185 cm, 190 cm or 195 cm. The fan tip radius may be around 110 cm, 115cm, 120 cm, 125 cm, 130 cm, 135 cm, 140 cm, 145 cm, 150 cm, 155 cm, 160cm, 165 cm, 170 cm, 175 cm, 180 cm, 185 cm, 190 cm or 195 cm. The fantip radius may be greater than 160 cm.

The fan tip radius may be in the range from 95 cm to 150 cm, optionallyin the range from 110 cm to 150 cm, optionally in the range of from 110cm to 145 cm, and further optionally in the range from 120 cm to 140 cm.

The fan tip radius may be in the range from 155 cm to 200 cm, optionallyin the range from 160 cm to 200 cm, and further optionally in the rangefrom 165 cm to 190 cm.

Optionally, the heat exchanger module has a flow area A_(HEX) and thefan module has a flow area A_(FAN), and a ratio of A_(FAN) to A_(HEX)being in the range of 0.3 to 0.8.

The flow area is to be understood to mean a cross-sectional area of theair flow taken perpendicularly to a central axis of the flow in the flowdirection. In other words, for the heat exchanger module the flow areaA_(HEX) corresponds to the cross-sectional area of the heat exchangermodule through which the flow passes. Likewise, for the fan assembly theflow area A_(FAN) corresponds to the cross-sectional area of the fanassembly through which the flow passes.

In one arrangement of the present disclosure, the flow area of the heatexchanger module has an annular profile and extends over only a radiallyoutward circumferential portion of the flow area of the fan assembly. Inother words, the air flow entering a radially proximal portion of theflow area of the fan assembly does not pass through the heat exchangerassembly and simply enters the fan assembly. In one arrangement, theradially outward circumferential portion of the flow area of the fanassembly amounts to 60% of the flow area of the fan assembly.

In another arrangement of the disclosure, the flow area of the heatexchanger module extends completely over the flow area of the fanassembly.

Optionally, the heat exchanger module has a fluid path diameter E,wherein the fluid path diameter E is greater than the fan diameter D.

In one embodiment, the heat exchanger module has a fluid path diameter Ethat is greater than the fan diameter D. In this embodiment, the inletduct that connects the heat exchanger module to the fan assembly has adiameter than converges from an exit from the heat exchanger module toan entrance to the fan assembly.

Optionally, the turbofan gas turbine engine further comprises an outerhousing, the outer housing enclosing the sequential arrangement of heatexchanger module, fan assembly, compressor module, and turbine module,an annular bypass duct being defined between the outer housing and thesequential arrangement of modules, a bypass ratio being defined as aratio of a mass air flow rate through the bypass duct to a mass air flowrate through the sequential arrangement of modules, and wherein thebypass ratio is less than 2.0.

A turbofan engine having a bypass ratio (BPR) of less than approximately2.0 will have a generally smaller bypass duct (the annular ductsurrounding the core gas turbine engine) than a turbofan engine having aBPR greater than approximately 2.0. For a turbofan engine with a BPRgreater than, say, 2.0, the correspondingly larger bypass duct volumeprovides more scope for positioning a heat exchanger within the bypassduct than would be the case for a low BPR turbofan engine.

Optionally, the fan assembly has two or more fan stages, at least one ofthe fan stages comprising a plurality of fan blades defining the fandiameter D.

In one arrangement, the fan assembly has two fan stages with both fanstages comprising a plurality of fan blades defining the same fandiameter. Alternatively, each of the fan stages may have different fandiameters.

Optionally, in use, an airflow entering the heat exchanger module with amean velocity of 0.4M, is divided between the set of vane airflowshaving a mean velocity of 0.2M, and the set of channel airflows having amean velocity of 0.6M.

In one arrangement of the present disclosure, the inlet and exhaustportions of the hollow vane may act as a diffuser to slow the air flowentering the hollow vane. In other words, the vane mass air flow isreduced from the mass air flow of the air flow entering the heatexchanger module. In this arrangement, the channel air flow through thechannels between circumferentially adjacent pairs of hollow vanes isincreased to maintain continuity of flow.

According to another aspect of the present disclosure there is provideda method of operating an aircraft comprising the gas turbine engineaccording to the first aspect, the method comprising taking off from arunway, wherein the maximum rotational speed of the turbine duringtake-off is in the range of from 8500 rpm to 12500 rpm.

The maximum take-off rotational fan speed may be in a range between 8500rpm to 12500 rpm. Optionally, for example for an engine with a fan tipradius in the range from 25 cm to 40 cm, the maximum take-off rotationalfan speed may be in a range between 9000 rpm to 11000 rpm. Optionally,for example for an engine with a fan tip radius in the range from 35 cmto 50 cm, the maximum take-off rotational fan speed may be in a rangebetween 8500 rpm to 10500 rpm.

According to another aspect of the present disclosure there is provideda method of operating a turbofan gas turbine engine, the gas turbineengine comprising, in axial flow sequence, a heat exchanger module, aninlet duct, a fan assembly, a compressor module, a turbine module, andan exhaust module, and wherein the method comprises the steps of:

(i) providing the fan assembly, the compressor module, the turbinemodule, and the exhaust module;

(ii) positioning the heat exchanger module in fluid communication withthe fan assembly by the inlet duct;

(iii) providing the heat exchanger module with a plurality of heattransfer elements for the transfer of heat from a first fluid containedwithin the heat transfer elements to an airflow passing over a surfaceof the or each heat transfer element prior to entry of the airflow intoan inlet to the fan assembly;

(iv) operating the gas turbine engine such that an airflow passing overa surface of the heat transfer element transfers heat energy from thefirst fluid contained within the heat transfer elements to the airflow;

(v) during operation of the gas turbine engine, monitoring anoperational parameter of the first fluid passing through the heattransfer elements and, if the operational parameter is outside of auser-defined acceptable parameter range, isolating the respective heattransfer element from the remaining heat transfer elements.

In one arrangement of the present disclosure, the first fluid is an oilthat is used to provide cooling to other parts of the turbofan engine.The oil is circulated through the heat transfer elements while the airflow passing over the heat transfer elements enables the rejection ofheat energy from the oil to the air flow.

Any physical damage to the heat transfer element may result in leakageof the oil from the heat transfer element. Such physical damage might becaused to the heat transfer element by a foreign object carried throughon the vane air flow.

Leakage of oil resulting from damage is undesirable since the oil isrequired for the heat rejection process. If more than a pre-determinedquantity of oil is lost from the oil system, the turbofan engine maybecome inoperative. Additionally, the downstream air from the gasturbine engine, for example a compressor bleed, may be used forventilation purposes and the presence of oil in such an air stream isundesirable.

The arrangement of the present disclosure enables a damaged heattransfer element to be isolated from the remaining heat transferelements. This will prevent further oil loss from the damaged elementand enable continued operation of the gas turbine engine albeit atperhaps slightly reduced heat energy rejection capability.

In one arrangement, a user may determine that a drop of, for example,20% in the first fluid pressure may be indicative of a leak of the firstfluid from one of the heat transfer elements. In other arrangements, adifferent degree of variation may be defined by the user.

Optionally, in step (v) the operational parameter of the first fluid isselected from the group comprising first fluid temperature, first fluidpressure, first fluid flow rate, and first fluid viscosity.

In one arrangement of the present disclosure, the flow rate of the firstfluid is monitored to determine whether damage has occurred to a heattransfer element. In alternative arrangements, another property of thefirst fluid could be used for this purpose, such as, for example, fluidpressure, fluid temperature, fluid viscosity or fluid dielectric.

According to another aspect of the present disclosure there is provideda turbofan gas turbine engine comprising, in axial flow sequence, a heatexchanger module, a fan assembly, a compressor module, a turbine module,and an exhaust module, the fan assembly comprising a plurality of fanblades defining a fan diameter (D), the heat exchanger module being influid communication with the fan assembly by an inlet duct, the heatexchanger module comprising a plurality of radially-extending hollowvanes arranged in a circumferential array with a channel extendingaxially between each pair of adjacent hollow vanes, an airflow enteringthe heat exchanger module being divided between a set of vane airflowsthrough each hollow vane, each vane airflow having a vane mass flow rateFlow_(Vane), and a set of channel airflows through each channel, eachchannel air flow having a channel mass flow rate Flow_(Chan), at leastone of the hollow vanes accommodating at least one heat transfer elementfor the transfer of heat from a first fluid contained within the or eachheat transfer element to the or each corresponding vane airflow passingover a surface of the or each heat transfer element;

and, in use, a Vane Airflow Ratio parameter V_(AR) is defined as:

${VAR} = \frac{Flow_{VaneTot}}{Flow_{ChanTot}}$

where: Flow_(VaneTot)=total mass flow rate of the vane mass flow rates,Flow_(Vane); and

-   -   Flow_(ChanTot)=total mass flow rate of the channel mass flow        rates, Flow_(Chan);        and the V_(AR) parameter is in the range of 0.05 to 3.0.

The VAR parameter provides a useful measure of the proportion of thetotal mass air flow that is entering the turbofan engine, and thus isingested by the fan assembly for combustion and power generation, thatpasses through the hollow vanes of the heat exchanger module and is thusavailable for heat energy rejection purposes.

The parameter of mass air flow rate (typically expressed in kg/s) iswell known to the skilled person, as is its measurement, and neitherwill be discussed further herein.

In an arrangement according to the present disclosure, between 5% and75% of the mass air flow entering the fan assembly will have passedthrough the hollow vanes and hence over a surface of the heat transferelements. This range has been determined to be sufficient to provide thecapability to reject operationally waste heat energy to the incoming airflow.

Optionally, the V_(AR) parameter is in the range of 0.43 to 1.0.

In an alternative arrangement of the present disclosure, the proportionof the mass air flow entering the fan assembly that has passed throughthe vanes and thus over a surface of the heat transfer elements isbetween 30% and 50%.

Optionally, the or each heat transfer element extends axially within thecorresponding hollow vane.

By extending axially along an interior volume of the hollow vane, theheat transfer element can efficiently transfer heat energy to theincoming vane air flow without the need to force the incoming vane airflow to change direction. This makes the heat exchanger module of thepresent disclosure more aerodynamically efficient and thus moreversatile and desirable for a user.

In alternative arrangements in which multiple heat transfer elements arepositioned within a hollow vane, these heat transfer elements may bepositioned side-by-side (i.e. circumferentially adjacent) or end-to end(i.e. axially adjacent). However, in each of these alternativearrangements, the individual heat transfer elements extend axiallywithin the corresponding hollow vane.

Optionally, the heat exchanger module has an annular flow area definedby an outer diameter and an inner diameter, each of the hollow vanesextends radially inwardly from the outer diameter and partially acrossthe annular flow area.

In an arrangement in which the required heat energy rejection capabilitycan be met by a heat exchanger module having a flow area that is lessthan the fan face area, then the annular depth of the heat exchangermodule may be reduced to provide a radially proximal annular region inwhich the incoming air flow bypasses the vanes and passes directly intothe fan assembly.

In this arrangement, the aerodynamic losses resulting from the incomingair flow passing through and between the hollow vanes can be reduced.This makes improves the efficiency of the gas turbine engine having thisalternative configuration.

According to another aspect of the present disclosure there is provideda method of operating a turbofan gas turbine engine, the gas turbineengine comprising, in axial flow sequence, a heat exchanger module, aninlet duct, a fan assembly, a compressor module, and a turbine module,and an exhaust module, and wherein the method comprises the steps of:

(i) providing the fan assembly, the compressor module, and the turbinemodule, and the exhaust module;

(ii) positioning the heat exchanger module in fluid communication withthe fan assembly by the inlet duct;

(iii) providing the heat exchanger module with a plurality ofradially-extending hollow vanes arranged in a circumferential array witha channel extending axially between each pair of adjacent hollow vanes,at least one of the hollow vanes accommodating at least one heattransfer element for the transfer of heat energy from a first fluidcontained within the or each heat transfer element to a correspondingvane airflow through the or each hollow vane and over a surface of theor each heat transfer element; and

(iv) operating the gas turbine engine such that an airflow entering theheat exchange module is divided between the set of vane airflows througheach hollow vane, each vane airflow having a vane mass flow rateFlow_(Vane), and a set of channel airflows through each channel, eachchannel airflow having a channel mass flow rate Flow_(Chan), with a VaneAirflow Ratio parameter V_(AR) being defined as:

${VAR} = \frac{Flow_{VaneTot}}{Flow_{C{hanTot}}}$

where: Flow_(VaneTot)=total mass flow rate of the vane airflows,Flow_(Vane); and

Flow_(ChanTot)=total mass flow rate of the channel airflows,Flow_(Chan);

and the V_(AR) parameter is in the range of 0.05 to 3.0.

In an arrangement according to the present disclosure, between 5% and75% of the mass air flow entering the fan assembly will have passedthrough the hollow vanes and hence over a surface of the heat transferelements. This range has been determined to be sufficient to provide thecapability to reject operationally waste heat energy to the incoming airflow.

Optionally, in step (iv) the engine is operated at a maximum dry thrustsetting, and the V_(AR) parameter is in the range of 0.43 to 1.0.

In an alternative arrangement of the present disclosure, the proportionof the mass air flow entering the fan assembly that has passed throughthe vanes and thus over a surface of the heat transfer elements isbetween 30% and 50%.

Optionally, in step (iii) the or each heat transfer element extendsaxially within the corresponding hollow vane.

By extending axially along an interior volume of the hollow vane, theheat transfer element can efficiently transfer heat energy to theincoming vane air flow without the need to force the incoming vane airflow to change direction. This makes the heat exchanger module of thepresent disclosure more aerodynamically efficient and thus moreversatile and desirable for a user.

In alternative arrangements in which multiple heat transfer elements arepositioned within a hollow vane, these heat transfer elements may bepositioned side-by-side (i.e. circumferentially adjacent) or end-to end(i.e. axially adjacent). However, in each of these alternativearrangements, the individual heat transfer elements extend axiallywithin the corresponding hollow vane.

According to another aspect of the present disclosure there is provideda turbofan gas turbine engine comprising, in axial flow sequence, a heatexchanger module, a fan assembly, a compressor module, a turbine module,and an exhaust module, the fan assembly comprising a plurality of fanblades defining a fan diameter (D), the heat exchanger module being influid communication with the fan assembly by an inlet duct, the heatexchanger module comprising a plurality of radially-extending hollowvanes arranged in a circumferential array, with a channel extendingaxially between each pair of adjacent hollow vanes, an airflow enteringthe heat exchange module being divided between a set of vane airflowsthrough each hollow vane, each vane airflow having a vane mass flow rateFlow_(Vane), and a set of channel airflows through each channel, eachchannel air flow having a channel mass flow rate Flow_(Chan), each ofthe hollow vanes accommodating at least one heat transfer element forthe transfer of heat from a first fluid contained within the or eachheat transfer element to the corresponding vane airflow passing over asurface of the or each heat transfer element;

wherein each hollow vane comprises, in axial flow sequence, an inletportion, a heat transfer portion, and an exhaust portion, the inletportion comprising a diffuser element, and the heat transfer portioncomprising at least one heat transfer element, and, the diffuser elementbeing configured to cause the vane mass flow rate Flow_(Vane) to belower than the channel mass flow rate Flow_(Chan) by a user-definedmargin.

The inclusion of a diffuser at the inlet to the hollow vane provides asmooth transition between the air flow upstream of the heat exchangermodule and the air flow passing into and through the hollow vane. Thediffuser slows the velocity of the air flow as the air flow enters thehollow vane and minimises the aerodynamic losses during this process.

Since it is the vane air flow passing over a surface of the heattransfer elements that enables the rejection of the heat energy to thevane air flow, a reduction in the velocity of the vane air flow improvesthe efficiency of the heat transfer process. This makes the heatexchanger module of the present disclosure more efficient at rejectingheat energy to the vane air flows.

Slowing the vane air flow passing through the hollow vane, relative tothe channel air flow passing through the channel between adjacent vanes,reduces the risk of damage that may be caused to the heat transferelements by foreign objects entering the hollow vane. This improves thereliability of the turbofan engine.

For example, in order for the vane air flows to provide for efficienttransfer of heat energy from the first fluid through the heat transferelements a user may require the vane mass flow rate Flow_(Vane) to beone third (⅓) of the channel mass flow rate Flow_(Chan). In alternativearrangements, this proportion may differ.

Optionally, the diffuser element comprises an axially extending firstduct, with an axial cross-section of the first duct having a divergentprofile in the axial flow direction.

In one arrangement of the present disclosure, the diffuser element takesthe form of a divergent passage extending downstream in the flowdirection from the inlet aperture of the hollow vane. In thisarrangement, the divergent profile is linear; in other words, in thediffuser element, the increase in the interior width of the hollow vane(in the circumferential direction) is directly proportional to the axialdistance along the diffuser element part of the hollow vane. Inalternative arrangements, this divergent profile may be non-linear andmay follow, for example, a parabolic or elliptical profile.

Optionally, the exhaust portion comprises an axially extending secondduct, with an axial cross-section of the second duct having a convergentprofile in the axial flow direction.

In another arrangement of the present disclosure, the hollow vane isprovided with a convergent profile at its exhaust portion. This featurehas the effect of speeding up the vane air flow passing exiting thehollow vane and thus enhances the effectiveness of the diffuserpositioned at the inlet to the hollow vane. For example, by speeding upthe air flow exiting the hollow vane it becomes easier to blend thisvane air flow with the faster moving channel air flows passing on eitherside of the vane air flow.

In the present embodiment, the convergent profile in the exhaust portionof the hollow vane has a linear profile. However, in other arrangementsthe convergent profile may also be non-linear; for example, logarithmicor elliptical.

Optionally, the or each heat transfer element extends axially within thecorresponding vane.

By extending axially along an interior volume of the hollow vane, theheat transfer element can efficiently transfer heat energy to theincoming vane air flow without the need to force the incoming vane airflow to change direction. This makes the heat exchanger module of thepresent disclosure more aerodynamically efficient and thus moreversatile and desirable for a user.

In alternative arrangements in which multiple heat transfer elements arepositioned within a hollow vane, these heat transfer elements may bepositioned side-by-side (i.e. circumferentially adjacent) or end-to end(i.e. axially adjacent). However, in each of these alternativearrangements, the individual heat transfer elements extend axiallywithin the corresponding hollow vane.

Optionally, one or more first vanes are positioned upstream of the inletportion such that air flow entering the hollow vane passes across the oreach first vane.

The use of one or more vanes upstream of the inlet to the hollow vanecan deflect foreign objects carried in the inlet air flow and that mightotherwise enter the hollow vane and cause mechanical damage to any heattransfer elements positioned inside the hollow vane. By blocking adirect line-of-sight into an inlet aperture of the hollow vane, theupstream vane(s) can prevent foreign objects from entering the hollowvane. This in turn can improve the robustness of the heat exchangermodule and consequently also improve the reliability of the heatexchanger module.

Optionally, one or more second vanes are positioned downstream of theexhaust portion such that air flow exhausted from the hollow vane passesacross the or each second vane.

By positioning one or more vanes at the exhaust outlet from the hollowvane it may be possible to further slow the vane air flow through thehollow vane, which improves the aerodynamic efficiency of the hollowvane.

As outlined above, the vane air flow exhausts the hollow vane at a lowervelocity than that of the channel air flow passing between vanes. Theappropriate positioning and orientation of vanes downstream of theexhaust portion of the hollow vane can be used to blend the lowervelocity vane air flow into the higher velocity channel air flow priorto the ingestion of the mixed flow by the fan assembly. Such blendingmay improve the aerodynamic efficiency of the fan assembly.

According to another aspect of the present disclosure there is provideda method of operating a turbofan gas turbine engine, the gas turbineengine comprising, in axial flow sequence, a heat exchanger module, aninlet duct, a fan assembly, a compressor module, a turbine module, andan exhaust module, and wherein the method comprises the steps of:

(i) providing the fan assembly, the compressor module, the turbinemodule, and the exhaust module;

(ii) positioning the heat exchanger module in fluid communication withthe fan assembly by the inlet duct;

(iii) providing the heat exchanger module with a plurality ofradially-extending hollow vanes arranged in a circumferential array witha channel extending axially through the heat exchanger module betweeneach pair of adjacent hollow vanes, such that an airflow entering theheat exchange module is divided between a set of vane airflows througheach of the hollow vanes, each vane airflow having a vane mass flow rateFlow_(Vane), and a set of channel airflows through each of the channels,each channel air flow having a channel mass flow rate Flow_(Chan);

(iv) providing each of the hollow vanes with, in axial flow sequence, aninlet portion, a heat transfer portion, and an exhaust portion, theinlet portion comprising a diffuser element, and the heat transferportion comprising at least one heat transfer element for the transferof heat energy from a first fluid contained within the or each heattransfer element to a corresponding vane airflow through the or eachhollow vane and over a surface of the or each heat transfer element; and

(v) sizing the diffuser elements such that when operating the gasturbine engine, a total vane mass flow rate Flow_(VaneTot) is lower thana total channel mass flow rate Flow_(ChanTot) by a user-defined margin.

The inclusion of a diffuser at the inlet to the hollow vane provides asmooth transition between the air flow upstream of the heat exchangermodule and the air flow passing into and through the hollow vane. Thediffuser slows the velocity of the air flow as the air flow enters thehollow vane and minimises the aerodynamic losses during this process.

Since it is the vane air flow passing over a surface of the heattransfer elements that enables the rejection of the heat energy to thevane air flow, a reduction in the velocity of the vane air flow improvesthe efficiency of the heat transfer process. This makes the heatexchanger module of the present disclosure more efficient at rejectingheat energy to the vane air flows.

Slowing the vane air flow passing through the hollow vane, relative tothe channel air flow passing through the channel between adjacent vanes,reduces the risk of damage that may be caused to the heat transferelements by foreign objects entering the hollow vane. This improves thereliability of the turbofan engine.

For example, in order for the vane air flows to provide for efficienttransfer of heat energy from the first fluid through the heat transferelements a user may require the vane mass flow rate Flow_(Vane) to beone third (⅓) of the channel mass flow rate Flow_(Chan). In alternativearrangements, this proportion may differ.

According to another aspect of the present disclosure there is provideda turbofan gas turbine engine comprising, in axial flow sequence, a heatexchanger module, a fan assembly, a compressor module, a turbine module,and an exhaust module, the fan assembly comprising a plurality of fanblades defining a fan diameter (D), the heat exchanger module being influid communication with the fan assembly by an inlet duct, the heatexchanger module further comprising a plurality of radially-extendinghollow vanes arranged in a circumferential array, with a channelextending axially between each pair of adjacent hollow vanes, each ofthe hollow vanes accommodating at least one heat transfer element forthe transfer of heat from a first fluid contained within the or eachheat transfer element to a corresponding vane airflow passing throughthe hollow vane and over a surface of the or each heat transfer element,each of the hollow vanes further comprising a flow modulator, the flowmodulator being configured to actively regulate the vane airflow as aproportion of a total airflow entering the heat exchanger module inresponse to a user requirement.

The active regulation of the vane airflow as a proportion of the totalairflow entering the heat exchanger module enables a user to determinethe proportion of the intake air flow that passes through the hollowvanes and therefore the proportion of the intake air that passes overthe heat transfer elements.

Since the vane airflow through the hollow vanes incurs an aerodynamicpenalty over the channel airflow that passes between adjacent vanes, byregulating the vane airflow to only that required to maintain a requiredlevel of heat energy rejection to the vane airflow, a user can maximisethe aerodynamic efficiency of the turbofan engine.

in one example configuration a user may require the vane air flow to be,for example, one third (⅓) of the channel air flow in order for the vaneair flows to provide for efficient transfer of heat energy from thefirst fluid through the heat transfer elements. In alternativeconfigurations, the proportion of the total air flow entering the heatexchange module that makes up the vane air flow may be greater or lessthan one third (⅓).

Optionally, the airflow entering the heat exchange module is dividedbetween the set of vane airflows through each hollow vane, each vaneairflow having a vane mass flow rate Flow_(Vane), and a set of channelairflows through each channel, each channel air flow having a channelmass flow rate Flow_(Chan), and the flow modulator is configured toactively regulate a ratio between a sum of the vane mass flow ratesFlow_(VaneTot) and a sum of the channel mass flow rates Flow_(ChanTot).

As outlined above, the heat energy rejected by the heat transferelements is transferred to the vane air flows passing through the hollowvanes and over the heat transfer elements. Specifically, the vane massflow rate is a controlling factor in the quantity of heat energytransferred to the vane air flows. Consequently, the ability for a userto regulate the proportion of the intake flow that passes through thehollow vanes thereby enables the user to control the quantity of heatenergy transferred to the vane air flows.

Optionally, each hollow vane comprises, in axial flow sequence, an inletportion, the at least one heat transfer element, and an exhaust portion,the inlet portion comprising a flow modulator, the flow modulator beingconfigured to restrict the vane airflow in response to the userrequirement.

In one arrangement of the present disclosure, the flow modulatorcomprises one or more first vanes positioned at the inlet portion. Thefirst vanes may be positioned upstream of an inlet aperture to thehollow vane. Alternatively, the flow modulator may comprise one or morefirst vanes that form the inlet portion of the hollow vane. In a furtheralternative arrangement, the first vanes may be positioned at the inletaperture to the hollow vane. The flow modulator in the form of the firstvanes can be actively regulated by a user to limit the vane air flowentering the hollow vane and thus to limit the vane mass flow rate.

Optionally, each hollow vane comprises, in axial flow sequence, an inletportion, the at least one heat transfer element, and an exhaust portion,the exhaust portion comprising a flow modulator, the flow modulatorbeing configured to restrict the vane airflow in response to the userrequirement.

In another arrangement of the present disclosure, the flow modulatorcomprises one or more second vanes positioned at the exhaust portion.Alternatively, the exhaust portion of the hollow vane may be formed bythe flow modulator in the form of the second vanes.

In yet another arrangement, the exhaust portion of the hollow vane maycomprise the flow modulator. For example, the flow modulator forming theexhaust portion may comprise a shape memory material that is configuredto restrict the vane airflow exhausted from the hollow vane in responseto the user requirement.

In each of these arrangements, the flow modulator in the form of thesecond vanes may be actively regulated by a user to restrict the vaneair flow exhausted from the hollow vane and in this way to limit thevane mass flow rate.

According to another aspect of the present disclosure there is provideda method of operating a turbofan gas turbine engine, the gas turbineengine comprising, in axial flow sequence, a heat exchanger module, aninlet duct, a fan assembly, a compressor module, a turbine module, andan exhaust module, and wherein the method comprises the steps of:

(i) providing the fan assembly, the compressor module, and the turbinemodule;

(ii) positioning the heat exchanger module in fluid communication withthe fan assembly by the inlet duct;

(iii) providing the heat exchanger module with a plurality ofradially-extending hollow vanes arranged in a circumferential array,with a channel extending axially through the heat exchanger modulebetween each pair of adjacent hollow vanes, such that an airflowentering the heat exchange module is divided between a set of vaneairflows through each of the hollow vanes, each vane airflow having avane mass flow rate Flow_(Vane), and a set of channel airflowsFlow_(Chan) through each of the channels, each channel air flow having achannel mass flow rate Flow_(Chan);

(iv) providing each of the hollow vanes with, at least one heat transferelement, and a flow modulator;

(v) operating the gas turbine engine including active control of theflow modulator to regulate a sum of the vane mass flow ratesFlow_(VaneTot) as a proportion of a total air mass flow entering theheat exchanger module in response to a user requirement.

The active regulation of the vane airflow as a proportion of the totalairflow entering the heat exchanger module enables a user to determinethe proportion of the intake air flow that passes through the hollowvanes and therefore the proportion of the intake air that passes overthe heat transfer elements.

Since the vane airflow through the hollow vanes incurs an aerodynamicpenalty over the channel airflow that passes between adjacent vanes, byregulating the vane airflow to only that required to maintain a requiredlevel of heat energy rejection to the vane airflow, a user can maximisethe aerodynamic efficiency of the turbofan engine.

in one example configuration a user may require the vane air flow to be,for example, one third (⅓) of the channel air flow in order for the vaneair flows to provide for efficient transfer of heat energy from thefirst fluid through the heat transfer elements. In alternativeconfigurations, the proportion of the total air flow entering the heatexchange module that makes up the vane air flow may be greater or lessthan one third (⅓).

The skilled person will appreciate that a feature described above inrelation to any one of the aspects may be applied, mutatis mutandis, toany other aspect of the invention. For example, in various embodimentsany two or more of the conditions for ratios as defined above, andoptionally all specified ratio ranges, may apply to any given aspect orembodiment. All aspects may apply to an engine of some embodiments.Furthermore, any feature described below may apply to any aspect and/ormay apply in combination with any one of the claims.

As noted elsewhere herein, the present disclosure may relate to aturbofan gas turbine engine. Such a gas turbine engine may comprise anengine core comprising a turbine, a combustor, a compressor, and a coreshaft connecting the turbine to the compressor. Such a gas turbineengine may comprise a fan (having fan blades) located upstream of theengine core. The fan may comprise any number of stages, for examplemultiple stages. Each fan stage may comprise a row of fan blades and arow of stator vanes. The stator vanes may be variable stator vanes (inthat their angle of incidence may be variable).

The turbofan gas turbine engine as described and/or claimed herein mayhave any suitable general architecture. For example, the gas turbineengine may have any desired number of shafts that connect turbines andcompressors, for example one, two or three shafts. Purely by way ofexample, the turbine connected to the core shaft may be a first turbine,the compressor connected to the core shaft may be a first compressor,and the core shaft may be a first core shaft. The engine core mayfurther comprise a second turbine, a second compressor, and a secondcore shaft connecting the second turbine to the second compressor. Thesecond turbine, second compressor, and second core shaft may be arrangedto rotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

In any turbofan gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of compressorstages, for example multiple stages. Each compressor stage may comprisea row of rotor blades and a row of stator vanes. The stator vanes may bevariable stator vanes (in that their angle of incidence may bevariable). The row of rotor blades and the row of stator vanes may beaxially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of turbine stages, for examplemultiple stages.

Each turbine stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.50, 0.49, 0.48, 0.47, 0.46,0.45, 0.44, 0.43, 0.42, 0.41, 0.40, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34,0.33, 0.32, 0.31, 0.30, 0.29, or 0.28. The ratio of the radius of thefan blade at the hub to the radius of the fan blade at the tip may be inan inclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds), for examplein the range of from 0.28 to 0.32. These ratios may commonly be referredto as the hub-to-tip ratio. The radius at the hub and the radius at thetip may both be measured at the leading edge (or axially forwardmost)part of the blade. The hub-to-tip ratio refers, of course, to thegas-washed portion of the fan blade, i.e. the portion radially outsideany platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 50 cm, 60 cm, 70 cm (around 27.5 inches), 80 cm(around 31.5 inches), 90 cm, 100 cm (around 39 inches), 110 cm (around43 inches), 120 cm (around 47 inches), 130 cm (around 51 inches), 140 cm(around 55 inches), 150 cm (around 59 inches), or 160 cm (around 130inches). The fan diameter may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 50 cm to 70 cmor 90 cm to 130 cm.

The fan face area may be equal to π multiplied by the square of the fantip radius.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 10000 rpm, for example less than 9000 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 50 cm to 90 cm (for example 60 cm to 80 cm or 65 cm to 75cm) may be in the range of from 7000 rpm to 10000 rpm, for example inthe range of from 7500 rpm to 10000 rpm, for example in the range offrom 8000 rpm to 9000 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 90 cm to 150 cm may bein the range of from 7000 rpm to 9000 rpm, for example in the range offrom 7500 rpm to 8600 rpm, for example in the range of from 8000 rpm to8600 rpm.

In use of the turbofan gas turbine engine, the fan (with associated fanblades) rotates about a rotational axis. This rotation results in thetip of the fan blade moving with a velocity U_(tip). The work done bythe fan blades 13 on the flow results in an enthalpy rise dH of theflow. A fan tip loading may be defined as dH/U_(tip) ², where dH is theenthalpy rise (for example the 1-D average enthalpy rise) across the fanand U_(tip) is the (translational) velocity of the fan tip, for exampleat the leading edge of the tip (which may be defined as fan tip radiusat leading edge multiplied by angular speed). The fan tip loading atcruise conditions may be greater than (or on the order of) any of: 0.22,0.23, 0.24, 0.25, 0.26, 0.27, 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34,0.35, 0.36, 0.37, 0.38, 0.39 or 0.40 (all values being dimensionless).The fan tip loading may be in an inclusive range bounded by any two ofthe values in the previous sentence (i.e. the values may form upper orlower bounds), for example in the range of from 0.28 to 0.31, or 0.29 to0.30.

Turbofan gas turbine engines in accordance with the present disclosuremay have any desired bypass ratio, where the bypass ratio is defined asthe ratio of the mass flow rate of the flow through the bypass duct tothe mass flow rate of the flow through the core at cruise conditions. Insome arrangements the bypass ratio may be greater than (or on the orderof) any of the following: 0.4, 0.5, 0.6, 0.7, 0.8, 0.9, 1.0, 1.1, 1.2,1.3, 1.4, or 1.5. The bypass ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of form of 0.4 to1.0, 0.5 to 0.9, or 0.6 to 0.9. The bypass duct may be substantiallyannular. The bypass duct may be radially outside the core engine. Theradially outer surface of the bypass duct may be defined by a nacelleand/or a fan case.

The overall pressure ratio of a turbofan gas turbine engine as describedand/or claimed herein may be defined as the ratio of the stagnationpressure upstream of the fan to the stagnation pressure at the exit ofthe highest-pressure compressor (before entry into the combustor). Byway of non-limitative example, the overall pressure ratio of a gasturbine engine as described and/or claimed herein at cruise may begreater than (or on the order of) any of the following: 10, 15, 20, 25,30, 35 or 40. The overall pressure ratio may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds), for example in the range of from20 to 35.

Specific thrust of a turbofan gas turbine engine may be defined as thenet thrust of the engine divided by the total mass flow through theengine. At cruise conditions, the specific thrust of an engine describedand/or claimed herein may be less than (or on the order of) any of thefollowing: 300 Nkg⁻¹ s, 350 Nkg⁻¹ s, 400 Nkg⁻¹ s, 450 Nkg⁻¹ s, 500 Nkg⁻¹s, 550 Nkg⁻¹ s, 600 Nkg⁻¹ s, 650 Nkg⁻¹ s or 700 Nkg⁻¹ s. The specificthrust may be in an inclusive range bounded by any two of the values inthe previous sentence (i.e. the values may form upper or lower bounds),for example in the range of from 300 Nkg⁻¹ s to 450 Nkg⁻¹ s, or 450Nkg⁻¹ s to 600 Nkg⁻¹ s. Such engines may be particularly efficient incomparison with conventional gas turbine engines.

A turbofan gas turbine engine as described and/or claimed herein mayhave any desired maximum thrust. Purely by way of non-limitativeexample, a gas turbine as described and/or claimed herein may be capableof producing a maximum thrust of at least (or on the order of) any ofthe following: 20 kN, 40 kN, 60 kN, 80 kN, 100 kN, 120 kN, 140 kN, 160kN, 180 kN, or 200 kN. The maximum thrust may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). Purely by way of example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust in the range of from 60 kN to 160 kN, for example 70 kNto 120 kN. The thrust referred to above may be the maximum net thrust atstandard atmospheric conditions at sea level plus 15 degrees C. (ambientpressure 101.3 kPa, temperature 30 degrees C.), with the engine static.

In use, the temperature of the flow at the entry to the high-pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K, 2000K, 2050K,2100K, 2150K, 2200K, 2250K or 2300K. The maximum TET may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds), for examplein the range of from 1800K to 2200K. The maximum TET may occur, forexample, at a high thrust condition, for example at a maximum take-off(MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example, at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium-based metal or an aluminium-based material(such as an aluminium-lithium alloy) or a steel-based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The turbofan gas turbine engines described and/or claimed herein may ormay not be provided with a variable area nozzle (VAN). Such a variablearea nozzle may allow the exit area of the bypass duct to be varied inuse. The general principles of the present disclosure may apply toengines with or without a VAN.

The fan of a turbofan gas turbine engine as described and/or claimedherein may have any desired number of fan blades, for example 12, 14,16, 18, 20, 22, 24 or 26 fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a giventurbofan gas turbine engine for an aircraft, the skilled person wouldimmediately recognise cruise conditions to mean the operating point ofthe engine at mid-cruise of a given mission (which may be referred to inthe industry as the “economic mission”) of an aircraft to which the gasturbine engine is designed to be attached. In this regard, mid-cruise isthe point in an aircraft flight cycle at which 50% of the total fuelthat is burned between top of climb and start of descent has been burned(which may be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of the gas turbine engine that provides athrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example, where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given turbofan gas turbine engine for an aircraft,cruise conditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.5 to 0.9, for example 0.55 to0.65, for example 0.75 to 0.85, for example 0.76 to 0.84, for example0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, forexample on the order of Mach 0.8, on the order of Mach 0.85 or in therange of from 0.8 to 0.85. Any single speed within these ranges may bepart of the cruise condition. For some aircraft, the cruise conditionsmay be outside these ranges, for example below Mach 0.7 or above Mach0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 40 kN to 65 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 70 kN to 95 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a turbofan gas turbine engine described and/or claimed hereinmay operate at the cruise conditions defined elsewhere herein. Suchcruise conditions may be determined by the cruise conditions (forexample the mid-cruise conditions) of an aircraft to which at least one(for example 2 or 4) gas turbine engine may be mounted in order toprovide propulsive thrust.

According to an aspect of the disclosure, there is provided an aircraftcomprising a turbofan gas turbine engine as described and/or claimedherein. The aircraft according to this aspect is the aircraft for whichthe gas turbine engine has been designed to be attached. Accordingly,the cruise conditions according to this aspect correspond to themid-cruise of the aircraft, as defined elsewhere herein.

According to an aspect of the disclosure, there is provided a method ofoperating a turbofan gas turbine engine as described and/or claimedherein. The operation may be at the cruise conditions as definedelsewhere herein (for example in terms of the thrust, atmosphericconditions and Mach Number).

According to an aspect of the disclosure, there is provided a method ofoperating an aircraft comprising a turbofan gas turbine engine asdescribed and/or claimed herein. The operation according to this aspectmay include (or may be) operation at the mid-cruise of the aircraft, asdefined elsewhere herein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Other aspects of the disclosure provide devices, methods and systemswhich include and/or implement some or all of the actions describedherein. The illustrative aspects of the disclosure are designed to solveone or more of the problems herein described and/or one or more otherproblems not discussed.

BRIEF DESCRIPTION OF THE DRAWINGS

There now follows a description of an embodiment of the disclosure, byway of non-limiting example, with reference being made to theaccompanying drawings in which:

FIG. 1 shows a schematic part-sectional view of a turbofan gas turbineengine according to the prior art;

FIG. 2 shows a schematic part-sectional view of a turbofan gas turbineengine according to a first embodiment of the disclosure;

FIG. 3 shows a schematic part-sectional view of a turbofan gas turbineengine according to a second embodiment of the disclosure;

FIG. 4 shows a perspective schematic view of the heat exchanger moduleof the turbofan engine of FIG. 2 showing the circumferential array ofvanes forming the heat exchanger module;

FIG. 5 shows a schematic part-sectional view of a part of the vane arrayof the heat exchanger module of FIG. 4;

FIG. 6 shows a schematic axial sectional view across one of the vanes ofthe vane array of the heat exchanger module of FIG. 4;

FIG. 7 shows a perspective schematic view of a heat transfer element ofthe heat exchanger module of FIG. 4;

FIG. 8A shows a schematic view of the vane of FIG. 6 with a flowmodulator in the form of first vanes arranged upstream of the inlet tothe vane;

FIG. 8B; shows a schematic view of the vane of FIG. 6 with a flowmodulator in the form of second vanes arranged downstream of the exhaustfrom the vane;

FIG. 8C shows a schematic view of the vane of FIG. 6 with a flowmodulator in the form of a VIGV array;

FIG. 9A shows a schematic axial sectional view across one of the vanesof the vane array of the heat exchanger module of FIG. 4 with the flowmodulator forming the exhaust portion of the vane, and the modulator ina ‘open’ position; and

FIG. 9B shows a schematic axial sectional view of the vane of FIG. 9with the modulator in a ‘closed’ position.

It is noted that the drawings may not be to scale. The drawings areintended to depict only typical aspects of the disclosure, and thereforeshould not be considered as limiting the scope of the disclosure. In thedrawings, like numbering represents like elements between the drawings.

DETAILED DESCRIPTION

FIG. 1 illustrates a conventional turbofan gas turbine engine 10 havinga principal rotational axis 9. The engine 10 comprises an air intake 12and a two-stage propulsive fan 13 that generates two airflows: a coreairflow A and a bypass airflow B. The gas turbine engine 10 comprises acore 11 that receives the core airflow A.

The engine core 11 comprises, in axial flow series, a low-pressurecompressor 14, a high-pressure compressor 15, combustion equipment 16, ahigh-pressure turbine 17, an intermediate-pressure turbine 18, alow-pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21surrounds the gas turbine engine 10 and defines a bypass duct 22 and abypass exhaust nozzle 18. The bypass airflow B flows through the bypassduct 22. The fan 13 is attached to and driven by the low-pressureturbine 19 via a shaft 26.

In use, the core airflow A is accelerated and compressed by thelow-pressure compressor 14 and directed into the high-pressurecompressor 15 where further compression takes place. The compressed airexhausted from the high-pressure compressor 15 is directed into thecombustion equipment 16 where it is mixed with fuel and the mixture iscombusted. The resultant hot combustion products then expand through,and thereby drive, the high-pressure, intermediate-pressure, andlow-pressure turbines 17, 18, 19 before being exhausted through thenozzle 20 to provide some propulsive thrust. The high-pressure turbine17 drives the high-pressure compressor 15 by a suitable interconnectingshaft 27. The low-pressure compressor 14 drives theintermediate-pressure turbine 18 via a shaft 28.

Note that the terms “low-pressure turbine” and “low-pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 13)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine. In some literature, the “low-pressure turbine” and“low-pressure compressor” referred to herein may alternatively be knownas the “intermediate-pressure turbine” and “intermediate-pressurecompressor”. Where such alternative nomenclature is used, the fan 13 maybe referred to as a first, or lowest pressure, compression stage.

Other turbofan gas turbine engines to which the present disclosure maybe applied may have alternative configurations. For example, suchengines may have an alternative number of fans and/or compressors and/orturbines and/or an alternative number of interconnecting shafts. By wayof further example, the gas turbine engine shown in FIG. 1 has a splitflow nozzle 20, 23 meaning that the flow through the bypass duct 22 hasits own nozzle 23 that is separate to and radially outside the coreengine nozzle 20. However, this is not limiting, and any aspect of thepresent disclosure may also apply to engines in which the flow throughthe bypass duct 22 and the flow through the core engine 11 are mixed, orcombined, before (or upstream of) a single nozzle, which may be referredto as a mixed flow nozzle. One or both nozzles (whether mixed or splitflow) may have a fixed or variable area. Whilst the described examplerelates to a turbofan engine, the disclosure may apply, for example, toany type of gas turbine engine, such as an open rotor (in which the fanstage is not surrounded by a nacelle) or turboprop engine, for example.

The geometry of the turbofan gas turbine engine 10, and componentsthereof, is defined by a conventional axis system, comprising an axialdirection (which is aligned with the rotational axis 9), a radialdirection (in the bottom-to-top direction in FIG. 1), and acircumferential direction (perpendicular to the page in the FIG. 1view). The axial, radial and circumferential directions are mutuallyperpendicular.

Referring to FIG. 2, a turbofan gas turbine engine according to a firstembodiment of the disclosure is designated generally by the referencenumeral 100. The turbofan gas turbine engine 100 comprises in axial flowsequence, a heat exchanger module 110, a fan assembly 130, a compressormodule 140, a turbine module 150, and an exhaust module 156.

In the present arrangement, the fan assembly 130 comprises two fanstages 131, with each fan stage 131 comprising a plurality of fan blades132. In the present arrangement each fan stage 131 has the same fandiameter 136, with the respective plurality of fan blades defining a fandiameter of 0.9 m. In an alternative arrangement, the two fan stages 131may have different fan diameters 136 each defined by the correspondingplurality of fan blades 132. As previously mentioned, the fan diameter(D) 136 is defined by a circle circumscribed by the leading edges of therespective plurality of fan blades 132.

FIG. 4 shows a perspective view of the heat exchanger module 110 and fanassembly 130 of the turbofan gas turbine engine 100 according to thefirst embodiment. The heat exchanger module 110 comprises twelveradially extending vanes 120 arranged in an equi-spaced circumferentialarray 122 with a channel 124 extending axially between each pair ofadjacent hollow vanes 120. Alternative embodiments may have more orfewer radially extending vanes 120.

Each of the vanes 120 is hollow and comprises four heat transferelements 112 arranged in a 2×2 configuration extending axially along thehollow interior of the vane 120, as shown in FIG. 6. Alternativeembodiments may not have a heat transfer element 112 within each vane120 or may have a different number of heat transfer elements 112 in anysingle vane 120.

Each of the heat transfer elements 112 has a corresponding swept area,which is the area of the heat transfer element 112 that is contacted bythe air flow 104. In the present arrangement, the total swept heattransfer element area (A_(HTE)) is the sum of the swept area of each ofthe individual heat transfer elements 112.

Each vane 120 is configured to allow the incoming airflow 104 passingthrough the heat exchange module 110 to pass through the hollow portionof the vane 120 and thence to flow over the respective heat transferelement 112. In this way heat energy is transferred from the first fluid190 to the air flow 104.

In use, as illustrated in FIG. 5, an airflow 104 entering the heatexchanger module 110 is divided between a set of vane airflows 106through each of the hollow vanes 120 and a set of channel airflows 108through each of the channels 124. Each of the vane airflows 106 has avane mass flowrate Flow_(Vane). Each of the channel airflows 108 has achannel mass flowrate Flow_(Chan).

A Vane Airflow Ratio parameter V_(AR) is defined as:

${VAR} = \frac{Flow_{VaneTot}}{Flow_{C{hanTot}}}$

where: Flow_(VaneTot)=total mass flow rate of the vane mass flow rates,Flow_(Vane); and

-   -   Flow_(ChanTot)=total mass flow rate of the channel mass flow        rates, Flow_(Chan)

In the present embodiment, the VAR parameter is 1.0. In other words, inthis arrangement the incoming airflow 104 is divided equally between thevane airflows 106 and the channel airflows 108.

FIG. 6 shows an axial cross-section through one of the hollow vanes 120and corresponds to the Section on ‘C-C’ from FIG. 5. Each of the hollowvanes 120 comprises, in axial flow sequence, an inlet portion 125, aheat transfer portion 126, and an exhaust portion 127.

The inlet portion 125 comprises a diffuser element 128. The diffuserelement 128 takes the form of an axially-extending first duct 128A. Thefirst duct 128A has an axial cross-section 1288 that has a linearlydivergent profile 128C. In use, the diffuser element 128 acts to slowthe incoming airflow 104 to the vane airflow 106. The diffuser element128 is sized such that the vane mass flow rate Flow_(Vane) is less thanthe channel mass flow rate Flow_(Chan) by a user-defined margin.

The heat transfer portion 126 accommodates the heat transfer elements112. Finally, the exhaust portion 127 comprises an axially-extendingsecond duct 127A having an axial cross-section 1278 that in turn has alinearly convergent profile 127C.

FIG. 7 shows a perspective schematic view of a radially outward facingsurface of a hollow vane 120. As outlined above, in this embodiment eachvane is provided with four heat transfer elements 112 arranged in a 2×2formation. Each of the heat transfer elements 112 has a fluid inlet 112Aand a corresponding fluid outlet 1128. The fluid inlet 112A provides afeed of hot oil (not shown) to the respective heat transfer element 112.Each of the fluid inlets 112A is provided with an actuatable inlet valve112C that can be switched to cut off the fluid flow into the heattransfer element 112. Additionally, each of the fluid outlets 1128 isprovided with an actuatable inlet valve 112D that can also be controlledto cut off the fluid flow leaving the heat transfer element 112.

Each of the fluid inlet valves 112C is provided with a fluid pressuresensor 112F that monitors the pressure of the oil flowing through theinlet 112A. In response to a sensed drop in the fluid pressure asmeasured by the pressure sensor 112F, the inlet valve 112C may beactuated to cut-off the oil flow through the respective heat transferelement 112. In this arrangement, the outlet valve 112D is also actuatedin response to a loss of oil pressure to thereby isolate thecorresponding heat transfer element 112 from the remaining oil flow.

Each of the fluid outlets 1128 is provided with a fluid flow sensor112E. In the event of a drop in fluid flow rate as detected by the flowsensor 112E, the corresponding fluid outlet valve 112D (and in thisarrangement, the corresponding fluid inlet valve 112C) can be actuatedto cut off the oil flow through the heat transfer element 112.

In addition to, or in an alternative to, the diffuser element 128described above, the hollow vane 120 may be provided with a flowmodulator 120A. The flow modulator 120A is configured to activelyregulate the vane airflow 106 as a proportion of a total airflow 104entering the heat exchanger module 110 in response to a userrequirement. In other words, the flow modulator 120A provides a userwith the ability to actively change the vane mass flowrateFlow_(VaneTot) as a proportion of the airflow 104 entering the turbofanengine.

In one arrangement, shown in FIG. 8A, the flow modulator 120A takes theform of first vanes 125A positioned upstream of the inlet portion 125 ofthe hollow vane 120. The first vanes 125A are actuatable to restrict thevane airflow 106 in response to the user requirement to change the ratioof the mass airflows between the hollow vane and the channel.

The first vanes 125A also provide a measure of protection to the heattransfer elements 112 positioned inside the hollow vane 120 from foreignobject damage caused by debris or other objects entering the hollow vane120.

FIG. 8B illustrates an alternative arrangement for the flow modulator120A in which second vanes 127D are positioned downstream of the exhaustportion 127 of the hollow vane 120. These second vanes 127D areactuatable to restrict the vane airflow 106 in response to the userrequirement to change the ratio of the mass airflows between the hollowvane 120 and the channel 124.

In a further alternative arrangement, FIG. 8C illustrates a flowmodulator 127E that takes the form of the exhaust portion 127 itself. Inthis arrangement, the heat exchange module 110 is close-coupled to anactuatable variable inlet guide vane (VIGV) array 127E. Theclose-coupling configuration provides for the fan assembly 130 toeffectively ‘suck’ the incoming air flow 104 through both the hollowvanes 120 and the channels 124. Such an arrangement also enables theinlet duct 160 to be shorter provides for an axially more compactturbofan gas turbine engine.

A further alternative form of the flow modulator 120A is illustrated inFIGS. 9A and 9B. In this arrangement the exhaust portion 127 of thehollow vane 120 is formed as the flow modulator 120A. In other words,the exhaust portion 127 is configured to change its shape in response tothe user requirement to change the ratio of the mass airflows betweenthe hollow vane 120 and the channel 124.

In the example shown in FIGS. 9A and 9B the exhaust portion 127 isformed from a shape memory alloy material. When the heat transferelements 112 are in use the temperature of the vane airflow 106 willincrease and the flow modulator 120A will open to thereby allowincreased vane mass airflow Flow_(Vane). Conversely when the heattransfer elements 112 are not in use the vane airflow 106 will have alower temperature, causing the flow modulator 120A to close and restrictthe vane mass airflow Flow_(Vane). This will therefore reduce theaerodynamic losses associated with flow through the hollow vanes 120when the heat transfer capability of the heat transfer elements 112 isnot required.

The heat exchanger module 110 is in fluid communication with the fanassembly 130 by an inlet duct 160. The heat exchange module 110 has anaxial length 115 of 0.4 m, this being 0.4 times the fan diameter of 0.9m.

The inlet duct 160 extends between a downstream-most face of the heattransfer elements and an upstream-most face of the fan assembly. In thepresent arrangement, the inlet duct 160 is linear. However, in otherarrangements the inlet duct 160 may be curved or convoluted.

The inlet duct 160 has a fluid path length 164 of 3.6 m, this being 4.0times the fan diameter of 0.9 m. The fluid path length 164 extends alonga central axis 162 of the inlet duct 160.

As outlined earlier, the heat exchanger module 110 has a flow area(A_(HEX)) 118. The heat exchanger module flow area 118 is thecross-sectional area of the heat exchanger module 110 through which anair flow 104 passes before being ingested by the fan assembly 130. Inthe present arrangement, the heat exchanger module flow area 118 has anannular cross-section and corresponds directly to the shape of the airflow passing through the heat exchanger module 110.

The fan assembly 130 has a corresponding flow area (A_(FAN)) 138. Thefan assembly flow area 138 is the cross-sectional area of the fanassembly 130 through which an air flow 104 passes before separating intoa core engine flow and a bypass flow.

The fan assembly flow area 138 has an annular shape since it correspondsto the annular area swept by the fan blades 132.

In the present arrangement, the heat exchanger module flow area 118 isequal to the fan assembly flow area 138, and the corresponding ratio ofA_(HEX)/A_(FAN) is equal to 1.0.

The heat exchanger module 110 has a flow diameter (E) 116, which is thediameter of the air flow passing through the heat exchanger module 110.In the present arrangement, the heat exchanger module flow diameter 116is equal to the fan diameter 136.

The heat exchanger module 110 comprises a plurality of heat transferelements 112 for the transfer of heat energy from a first fluid 190contained within the heat transfer elements 112 to an airflow 104passing over a surface 113 of the heat transfer elements 112 prior toentry of the airflow 104 into the fan assembly 130. In the presentembodiment, the first fluid 190 is a mineral oil. In other arrangements,the first fluid 190 may be an alternative heat transfer fluid such as,for example, a water-based fluid, or the fuel used by the turbofan gasturbine engine.

The heat transfer elements 112 have a conventional tube and finconstruction and will not be described further. In an alternativearrangement, the heat transfer elements may have a differentconstruction such as, for example, plate and shell.

The turbofan gas turbine engine 100 further comprises an outer housing170. The outer housing 170 fully encloses the sequential arrangement ofthe heat exchanger module 110, inlet duct 160, fan assembly 130,compressor module 140, and turbine module 150. The outer housing 170defines a bypass duct 180 between the outer housing 170 and the coreengine components (comprising inter alia the compressor module 140 andthe turbine module 150). In the present arrangement, the bypass duct 180has a generally axi-symmetrical annular cross-section extending over thecore engine components. In other arrangements, the bypass duct 180 mayhave a non-symmetric annular cross-section or may not extend around acomplete circumference of the core engine components.

Referring to FIG. 3, a turbofan gas turbine engine according to a secondembodiment of the disclosure is designated generally by the referencenumeral 200. Features of the turbofan gas turbine engine 200 whichcorrespond to those of turbofan gas turbine engine 100 have been givencorresponding reference numerals for ease of reference.

The turbofan gas turbine engine 200 comprises in axial flow sequence, aheat exchanger module 210, a fan assembly 130, a compressor module 140,and a turbine module 150.

The fan assembly 130, compressor module 140, and turbine module 150correspond directly to the those of the first embodiment describedabove.

The heat exchanger module 210 comprises a plurality of heat transferelements 212 and is also in fluid communication with the fan assembly130 by an inlet duct 260. As in the first embodiment, the inlet duct 260extends between a downstream-most face of the heat transfer elements andan upstream-most face of the fan assembly.

The inlet duct 260 has a fluid path length 264 along a central axis 162of the inlet duct 260 of 2.4 m, this being 2.7 times the fan diameter of0.9 m.

The heat exchanger module 210 has a flow area (A_(HEX)) 218. As in thefirst embodiment, the heat exchanger module flow area 118 is annular incross-section.

However, in this arrangement the heat transfer elements 212 do notextend completely across that cross-section of the heat exchange module210 that is available for the flow 104. In other words, there is aradially proximal portion of the cross-section of the heat transfermodule across which there are no heat transfer elements 212.

The fan assembly 130 has a flow area (A_(FAN)) 138 that, as describedabove, has an annular shape corresponding to the annular area swept bythe fan blades 132.

In the present arrangement, despite the heat exchanger module flow area218 having different dimensions to the fan assembly flow area 138, theheat exchanger module flow area 218 is equal to the fan assembly flowarea 138. As for the first embodiment, the corresponding ratio ofA_(HEX)/A_(FAN) is equal to 1.0.

The heat exchanger module 210 has a flow diameter 216. The heatexchanger module flow diameter 216 is greater than the fan diameter 136.

The turbofan gas turbine engine 200 further comprises an outer housing270. As with the first embodiment described above, the outer housing 170fully encloses the sequential arrangement of the heat exchanger module210, inlet duct 260, fan assembly 130, compressor module 140, andturbine module 150. The outer housing 270 also defines an annular bypassduct 180 between the outer housing 170 and the core engine components

In use the turbofan gas turbine engine 200 functions in the same manneras described above in relation to the turbofan gas turbine engine 100 ofthe first embodiment.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

The invention includes methods that may be performed using the subjectdevices. The methods may comprise the act of providing such a suitabledevice. Such provision may be performed by the end user. In other words,the “providing” act merely requires the end user obtain, access,approach, position, set-up, activate, power-up or otherwise act toprovide the requisite device in the subject method. Methods recitedherein may be carried out in any order of the recited events which islogically possible, as well as in the recited order of events.

In addition, where a range of values is provided, it is understood thatevery intervening value, between the upper and lower limit of that rangeand any other stated or intervening value in that stated range, isencompassed within the invention.

Except where mutually exclusive, any of the features may be employedseparately or in combination with any other features and the disclosureextends to and includes all combinations and sub-combinations of one ormore features described herein.

1. A turbofan gas turbine engine comprising, in axial flow sequence, aheat exchanger module, a fan assembly, a compressor module, a turbinemodule, and an exhaust module, the fan assembly comprising a pluralityof fan blades defining a fan diameter, the heat exchanger module beingin fluid communication with the fan assembly by an inlet duct, the heatexchanger module comprising a plurality of heat transfer elements fortransfer of heat from a first fluid contained within the heat transferelements to an airflow passing over a surface of the heat transferelements prior to entry of the airflow into an inlet to the fanassembly, wherein each heat transfer element may be individually andindependently fluidly isolated from the remaining heat transferelements.
 2. The turbofan gas turbine engine as claimed in claim 1,wherein each heat transfer element comprises an inlet and an outlet, andboth the inlet and the outlet are selectably closable to fluidly isolatethe heat transfer element.
 3. The turbofan gas turbine engine as claimedin claim 1, wherein the heat exchanger module further comprises a flowsensor configured to detect a flow rate of the first fluid through theheat transfer elements, and the heat exchanger module is configured toisolate a heat transfer element in the event that the detected flow rateis outside of a predetermined acceptable flow rate range.
 4. Theturbofan gas turbine engine as claimed in claim 1, wherein the heatexchanger module further comprises a pressure sensor configured todetect a pressure of the first fluid contained within the heat transferelements, and the heat exchanger module is configured to isolate a heattransfer element in the event that the detected pressure is outside of apredetermined acceptable pressure range.
 5. The turbofan gas turbineengine as claimed in claim 1, wherein the heat exchanger modulecomprises a plurality of radially-extending hollow vanes arranged in acircumferential array with a channel extending axially through the heatexchanger module between each two circumferentially adjacent vanes, eachof the hollow vanes accommodating at least one heat transfer element. 6.The turbofan gas turbine engine as claimed in claim 1, wherein the fandiameter is within the range of 0.3 m to 2.0 m.
 7. The turbofan gasturbine engine as claimed in claim 1, wherein the heat exchanger modulehas a flow area A_(HEX) and the fan module has a flow area A_(FAN), anda ratio of A_(FAN) to A_(HEX) being in the range of 0.3 to 0.8.
 8. Theturbofan gas turbine engine as claimed in claim 1, wherein the heatexchanger module has a fluid path diameter, wherein the fluid pathdiameter is greater than the fan diameter.
 9. The turbofan gas turbineengine as claimed in claim 1, the turbofan gas turbine engine furthercomprising an outer housing, the outer housing enclosing the sequentialarrangement of heat exchanger module, fan assembly, compressor module,and turbine module, an annular bypass duct being defined between theouter housing and the sequential arrangement of modules, a bypass ratiobeing defined as a ratio of a mass air flow rate through the bypass ductto a mass air flow rate through the sequential arrangement of modules,and wherein the bypass ratio is less than 2.0.
 10. The turbofan gasturbine engine as claimed in claim 1, wherein the fan assembly has twoor more fan stages, at least one of the fan stages comprising aplurality of fan blades defining the fan diameter.
 11. The turbofan gasturbine engine as claimed in claim 1, wherein, in use, an airflowentering the heat exchanger module with a mean velocity of 0.4M, isdivided between a first airflow through the hollow vanes having a meanvelocity of 0.2M, and a second airflow through the channels betweenadjoining pairs of hollow vanes having a mean velocity of 0.6M.
 12. Amethod of operating an aircraft comprising the gas turbine engine asclaimed in claim 1, the method comprising taking off from a runway,wherein the maximum rotational speed of the turbine during take-off isin the range of from 8500 rpm to 12500 rpm.
 13. A method of operating aturbofan gas turbine engine, the gas turbine engine comprising, in axialflow sequence, a heat exchanger module, an inlet duct, a fan assembly, acompressor module, a turbine module, and an exhaust module, and whereinthe method comprises the steps of: (i) providing the fan assembly, thecompressor module, the turbine module, and the exhaust module; (ii)positioning the heat exchanger module in fluid communication with thefan assembly by the inlet duct; (iii) providing the heat exchangermodule with a plurality of heat transfer elements for the transfer ofheat from a first fluid contained within the heat transfer elements toan airflow passing over a surface of each heat transfer element prior toentry of the airflow into an inlet to the fan assembly; (iv) operatingthe gas turbine engine such that an airflow passing over a surface ofthe heat transfer element transfers heat energy from the first fluidcontained within the heat transfer elements to the airflow; (v) duringoperation of the gas turbine engine, monitoring an operational parameterof the first fluid passing through the heat transfer elements and, ifthe operational parameter is outside of a user-defined acceptableparameter range, isolating the respective heat transfer element from theremaining heat transfer elements.
 14. The method as claimed in claim 13,wherein in step (v) the operational parameter of the first fluid isselected from the group comprising first fluid temperature, first fluidpressure, first fluid flow rate, first fluid viscosity, and first fluiddielectric.